Gas turbine engine

ABSTRACT

A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine has high efficiency together with low noise, in particular the noise emanating from the front of the fan. The contribution of the fan noise emanating from the front of the engine to the Effective Perceived Noise Level (EPNL) at a take-off lateral reference point, defined as the point on a line parallel to and 450 m from the runway centre line where the EPNL is a maximum during take-off, is in the range of from 0 EPNdB and 12 EPNdB lower than the contribution of the fan noise emanating from the rear of the engine to the EPNL at the take-off lateral reference point.

The present disclosure relates to a gas turbine engine having animproved noise signature.

Gas turbine engines are typically optimized to provide high efficiency,because this generally results in lower fuel burn, and thus lowerrunning costs. However, the noise generated by a gas turbine engine usedto power an aircraft is an important factor due to the impact thataircraft noise can have on communities.

In this regard, gas turbine engines generate a significant proportion ofthe noise produced by an aircraft. Regulations define an “EffectivePerceived Noise Level” (EPNL) which is a measure of the impact of thegenerated noise as perceived by the human ear, taking into accountfactors such as frequency, absolute level, tonal components and durationof the noise.

A turbofan gas turbine engine comprises a number of different noisesources. For example, the fan itself is a source of noise, and that fannoise can be separated into two distinct components: a component thatemanates in a forwards direction from the front of the engine; and acomponent that emanates in a rearward direction from the rear of theengine. Further noise sources include (but are not limited to) noisefrom the jet stream exhausted from the engine, noise from the turbine atthe rear of the engine, and noise from the installation of the engine onthe aircraft.

It is desirable to reduce the perceived noise of a gas turbine engine soas to reduce the impact of the noise on the human ear.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and

a gearbox that receives an input from the core shaft (26) and outputsdrive to the fan so as to drive the fan at a lower rotational speed thanthe core shaft, wherein:

the contribution of the fan noise emanating from the front of the engineto the Effective Perceived Noise Level (EPNL) at a take-off lateralreference point, defined as the point on a line parallel to and 450 mfrom the runway centre line where the EPNL is a maximum during take-off,is in the range of from 0 EPNdB and 12 EPNdB, optionally 0 EPNdB and 10EPNdB, lower than the contribution of the fan noise emanating from therear of the engine to the EPNL at the take-off lateral reference point.

According to an aspect, there is provided a method of operating a gasturbine engine attached to an aircraft, wherein the gas turbine enginecomprises:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox that receives an input from the core shaft (26) and        outputs drive to the fan so as to drive the fan at a lower        rotational speed than the core shaft, and wherein

the method comprises using the gas turbine engine to provide thrust tothe aircraft for taking off from a runway, during which:

the contribution of the fan noise emanating from the front of the engineto the Effective Perceived Noise Level (EPNL) at a take-off lateralreference point, defined as the point on a line parallel to and 450 mfrom the runway centre line where the EPNL is a maximum during take-off,is in the range of from 0 EPNdB and 12 EPNdB, optionally 0 EPNdB and 10EPNdB, lower than the contribution of the fan noise emanating from therear of the engine to the EPNL at the take-off lateral reference point.

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine, wherein the gas turbine enginecomprises:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox that receives an input from the core shaft and outputs        drive to the fan so as to drive the fan at a lower rotational        speed than the core shaft, and wherein

the method comprises taking off from a runway, during which:

the contribution of the fan noise emanating from the front of the engineto the Effective Perceived Noise Level (EPNL) at a take-off lateralreference point, defined as the point on a line parallel to and 450 mfrom the runway centre line where the EPNL is a maximum during take-off,is in the range of from 0 EPNdB and 12 EPNdB, optionally 0 EPNdB and 10EPNdB, lower than the contribution of the fan noise emanating from therear of the engine to the EPNL at the take-off lateral reference point.

According to an aspect, there is provided a gas turbine enginecomprising at least one (for example 2 or 4) gas turbine engine asdescribed and/or claimed herein.

As referred to herein, including in the claims, the Effective PerceivedNoise Level (EPNL) is as calculated in the conventional manner, asdefined in Appendix 2 of the Fifth Edition (July 2008) of Annex 16(Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise). For completeness, the calculationof the EPNL from measured noise data is as defined in Section 4 ofAppendix 2 of the Fifth Edition (July 2008) of Annex 16 (EnvironmentalProtection) to the Convention on International Civil Aviation, Volume 1(Aircraft Noise), from page APP 2-13 to APP 2-21. For completeness, theEPNL is defined at the reference atmospheric conditions provided inSection 3.6.1.5 of the Fifth Edition (July 2008) of Annex 16(Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise).

Also as referred to herein, the take-off lateral reference point isdefined as the point on a line parallel to and 450 m from the runwaycentre line where the EPNL is a maximum during take-off, as defined inSection 3.3.1, a), 1) of the Fifth Edition (July 2008) of Annex 16(Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise).

Conventionally, the relative contribution of the fan noise emanatingfrom the front of the engine at the take-off lateral reference pointwould be greater, and in some cases may be higher than the contributionof the fan noise emanating from the rear of the engine (also referred toherein as “rear fan noise”) to the EPNL at the take-off lateralreference point. Accordingly, for a given level of rear fan noise, gasturbine engines in accordance with the present disclosure have a lowercombined fan front and rear EPNL at the take-off lateral referencepoint, and thus have a lower noise impact on communities in the vicinityof the runway.

The present inventors have realised that for a given power of gasturbine engine, the use of a gearbox between the fan and the turbinethat drives the fan enables the fan forward noise to be reduced relativeto the rear fan noise. This may be through using the relatively lowrotational speed of the fan (due to the gearbox) for a given enginethrust level to reduce the fan tip Mach number.

According to any aspect, the relative Mach number of the tip of each fanblade may be in the range of from 0.8 M to 1.09 M, optionally 0.9 M to1.08 M, optionally 1.0 M to 1.07 M, optionally less than 1.09 M at thetake-off lateral reference point.

Such a relative Mach number—which is lower than conventional designs—mayhelp to reduce the noise generated by the fan. In particular, havingsuch a fan tip relative Mach number at the take-off lateral referencepoint ensures that the supersonic noise generated by the fan is kept toan acceptable level, and reduces the level of fan noise propagatingforwards of the engine relative to conventional designs. As anadditional or alternative benefit, this, in turn, may reduce the amountof acoustic liner required on the intake of the engine, which mayfacilitate a shorter intake (albeit acoustic liner may be provided to asignificant proportion of the intake). For gas turbine engines havingthe fan diameters referred to herein (which are larger than manyprevious engines), the intake may contribute significantly to theaerodynamic drag on the engine during use, and so the ability to reduceits extent may be particularly beneficial.

As used herein, the relative Mach number at the tip of the fan blade maybe defined as the vector sum of the axial Mach number due to the forwardspeed of the engine and the rotational Mach number due to the rotationof the fan blades about the engine axis.

The relative Mach number of the tip of each fan blade may not exceed1.09 M during take-off of an aircraft to which the gas turbine engine isattached. For example, the relative Mach number of the tip of each fanblade may have a maximum value in the range of from 0.8 M to 1.09 M,optionally 0.9 M to 1.08 M, optionally 1.0 M to 1.07 M during take-offof an aircraft to which the gas turbine engine is attached. In thisregard, the take-off may be considered to last at least as long asnecessary to determine the maximum point of the EPNL between release ofbrake and top of climb of the aircraft. In practice, this is likely tobe within a horizontal distance of 10 km or less of the release ofbrake.

A bypass duct may be defined between an inner flow boundary formed bythe engine core and an outer flow boundary formed by a nacelle. Anintake may be defined as the radially outer flow boundary of the flowinto the engine upstream of the tips of the leading edge of the fanblades.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades;

a gearbox that receives an input from the core shaft and outputs driveto the fan so as to drive the fan at a lower rotational speed than thecore shaft;

a bypass duct defined between an inner flow boundary formed by theengine core and an outer flow boundary formed by a nacelle; and

an intake defined as the radially outer flow boundary of the flow intothe engine upstream of the tips of the leading edge of the fan blades,wherein:

-   -   a bypass noise attenuation proportion L is defined as:

$L = \frac{J + H}{2G}$

where:

G is the axial length between the tip of the trailing edges of the fanblades and the trailing edge of the nacelle;

H is the total axial length of acoustic attenuation material (320)provided to the outer flow boundary of the bypass duct over the axialextent between the tip of the trailing edges of the fan blades and thetrailing edge of the nacelle; and

J is the total axial length of acoustic attenuation material (330)provided to the inner flow boundary of the bypass duct over the axialextent between the tip of the trailing edges of the fan blades and thetrailing edge of the nacelle;

-   -   an intake noise attenuation proportion k is defined as:

$K = \frac{E}{F}$

where:

E is the total axial length of acoustic attenuation material (310)provided to the intake; and

F is the axial length of the intake; and

a forward to rearward noise attenuation proportion M is in the range offrom 0.8 to 2.5, optionally 1.1 to 2.3, optionally 1.2 to 2.1,optionally 1.3 to 2, optionally 1.4 to 1.8, optionally on the order of1.6, where:

$M = \frac{K}{L}$

Indeed, according to any aspect, a bypass noise attenuation proportion Lmay be defined as:

$L = \frac{J + H}{2G}$

where:

G is the axial length between the tip of the trailing edges of the fanblades and the trailing edge of the nacelle;

H is the total axial length of acoustic attenuation material provided tothe outer flow boundary of the bypass duct over the axial extent betweenthe tip of the trailing edges of the fan blades and the trailing edge ofthe nacelle; and

J is the total axial length of acoustic attenuation material provided tothe inner flow boundary of the bypass duct over the axial extent betweenthe tip of the trailing edges of the fan blades and the trailing edge ofthe nacelle.

An intake noise attenuation proportion k may be defined as:

$K = \frac{E}{F}$

where:

E is the total axial length of acoustic attenuation material provided tothe intake; and

F is the axial length of the intake.

A forward to rearward noise attenuation proportion M may be in the rangeof from 0.8 to 2.5, optionally 1.1 to 2.3, optionally 1.2 to 2.1,optionally 1.3 to 2, optionally 1.4 to 1.8, optionally on the order of1.6, where:

$M = \frac{K}{L}$

The bypass noise attenuation proportion L may be in the range of from0.4 to 0.7, optionally, 0.45 to 0.65, optionally 0.5 to 0.6.

The intake noise attenuation proportion K may be in the range of from0.55 to 0.95, optionally, 0.6 to 0.9, optionally 0.7 to 0.8.

Acoustic attenuation material may be material whose primary, or evensole, function is to attenuate noise. Such material may not be providedto the engine but for its acoustic attenuation properties. Such materialmay comprise holes which are open to the main flow (i.e. the flowcontaining the noise to be attenuated) on one side. Such holes may openon their other side to cavities. Thus, the acoustic attenuation materialmay comprise holes that fluidly connect the main flow (e.g. the bypassflow or intake flow) with cavities. The number of cavities may or maynot be equal to the number of holes in such an arrangement.

Providing a forward to rearward noise attenuation proportion M and/or abypass noise attenuation proportion L and/or an intake noise attenuationproportion K in the ranges described and/or claimed herein may result ina gas turbine engine having acceptable fan noise levels, and havingimproved installation properties, such as acceptably low drag (forexample through a shorter than conventional intake section).

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at thetake-off lateral reference point may be less than 2800 rpm, for exampleless than 2500 rpm, for example less than 2300 rpm. Purely by way offurther non-limitative example, the rotational speed of the fan at thetake-off lateral reference point for an engine having a fan diameter inthe range of from 220 cm to 290 cm (for example 230 cm to 270 cm) may bein the range of from 1700 rpm to 2800 rpm, for example in the range offrom 2000 rpm to 2600 rpm, for example in the range of from 2000 rpm to2500 rpm. Purely by way of further non-limitative example, therotational speed of the fan at take-off lateral reference point for anengine having a fan diameter in the range of from 320 cm to 400 cm (forexample 330 cm and 370 cm) may be in the range of from 1200 rpm to 2000rpm, for example in the range of from 1300 rpm to 1800 rpm, for examplein the range of from 1400 rpm to 1600 rpm.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 and 20. For example thebypass ratio may be in the range of from 12.5 to 18, for example 13 to18, for example 13 to 17. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass duct may besubstantially annular. The bypass duct may be radially outside the coreengine. The radially outer surface of the bypass duct may be defined bya nacelle and/or a fan case.

According to any aspect, the ratio of the mean total pressure of theflow at the fan exit that subsequently flows through the bypass duct tothe mean total pressure of the flow at the fan inlet, may be in therange of from 1.25 to 1.5, optionally 1.35 to 1.45, at the take-offlateral reference point. This optional feature may, for somearrangements, help to facilitate the average velocity of the flow at theexit to the bypass duct and/or reduced jet noise.

As noted elsewhere herein, the gas turbine engine may further comprisean intake that extends upstream of the fan blades. A radially innerintake length L may be defined as the axial distance between the leadingedge of the intake and the leading edge of the fan blades at the hub.The fan diameter D may be the diameter of the fan at the leading edge ofthe tips of the fan blades. The ratio L/D may be in the range of from0.2 to 0.5, optionally 0.25 to 0.5, optionally 0.3 to 0.5, optionally0.32 to 0.49, optionally 0.33 to 0.48.

As noted elsewhere herein, a gas turbine engine has multiple noisesources, only one of which is the noise generated by the fan. Forexample, for a conventional engine, the noise generated by the jet alsocontributes significantly to the overall noise of the engine.Accordingly, in some cases, it may be further advantageous toadditionally reduce the jet noise

In some arrangements, the contribution of the jet to the EffectivePerceived Noise Level (EPNL) at the take-off lateral reference point maybe in the range of from 0 EPNdB and 15 EPNdB, optionally 2 dB and 12 dB,lower than the contribution of the fan noise emanating from the rear ofthe engine to the EPNL at the take-off lateral reference point.

The jet noise may comprise noise generated by the flow exiting the coreof the engine and the flow exiting a bypass duct. As referred to herein,the bypass duct may be defined radially outside the engine core and aradially inside a nacelle. The majority of the jet noise may begenerated by the flow exiting the bypass duct.

The average velocity of the flow at the exit to the bypass duct may bein the range of from 200 m/s to 275 m/s, 200 m/s to 270 m/s, for example200 m/s to 265 m/s, for example 230 m/s to 265 m/s at the take-offlateral reference point.

Controlling the average velocity of the flow at the exit to the bypassduct to be within the above ranges may at least in part facilitate jetnoise reduction, whilst retaining high propulsive efficiency. Withoutbeing bound by any particular theory, this may be due to a reduction inthe strength of the shear layer between the bypass flow and thesurrounding air as the bypass flow exits the engine.

The average velocity of the flow at the exit to the bypass duct at atake-off lateral reference point may be in the range of from 50 m/s to90 m/s lower, optionally 55 m/s to 85 m/s lower, optionally 60 m/s to 85m/s lower, than the average velocity of the flow at the exit to thebypass duct at cruise conditions.

Providing and/or operating an engine such that the average velocity ofthe flow at the exit to the bypass duct at a take-off lateral referencepoint is in the above ranges lower than at cruise may mean that thenoise generated at the take-off lateral reference point is reduced. Thismay be because of the reduction in the strength of the shear layerbetween the bypass flow and the surrounding air as the bypass flow exitsthe engine. In conventional engines, the magnitude of this differencemay be lower. In terms of the noise of an aircraft, this may beparticularly significant, because the noise generated at take-off, whenthe aircraft is close to the ground and potentially close tocommunities, can have more of a significant impact than the noisegenerated at cruise.

The noise generated by the turbine may also typically contributesignificantly to the overall noise of the engine. Accordingly, in somecases, it may be further advantageous to additionally reduce the turbinenoise. In this regard, for a given power of gas turbine engine, the useof a gearbox between the fan and the turbine that drives the fan mayalso enable the turbine noise to be reduced, if desired. This may bebecause the increased rotational speed of the turbine (relative to thefan) allows the frequency of at least some of the tones generated by theturbine (at least some of which may be referred to as fundamental bladepassing frequencies) to be increased. Such increased frequencies mayalso be subject to an increase in atmospheric attenuation relative toconventional turbine frequencies. As such, these tones are less wellperceived by the human ear (and possibly at least some tones are not atall perceived by the human ear), and so are given a lower weighting inthe EPNL calculation (even a zero weighting if the frequency is highenough, i.e. not included), thereby reducing the contribution of theturbine noise to the EPNL relative to the rear fan noise.

The turbine that drives the fan via the gearbox may comprise at leasttwo axially separated rotor stages. For example, the turbine that drivesthe fan via the gearbox may comprise two, three, four, five or greaterthan five axially separated rotor stages. A rotor stage may be part of aturbine stage that also comprises a stator vane stage, which may beaxially separated from the respective rotor stage of the turbine stage.Each rotor stage of the turbine that drives the fan via the gearbox maybe separated from at least one immediately upstream and/or downstreamrotor stage by a row of stator vanes.

Optionally, the fan diameter may be in the range of from 320 cm to 400cm, and the turbine that drives the fan via the gearbox may comprise 4stages. Further optionally, the fan diameter may be in the range of from220 cm to 290 cm, and the turbine that drives the fan via the gearboxmay comprise 3 stages.

The number of turbine blades in the rotor stages of the turbine thatdrives the fan via the gearbox may influence the frequency of at leastsome of the tones generated by the turbine, and thus may assist inallowing the fundamental frequency or frequencies to be moved to a rangewhere they are less well perceived by the human ear (and possibly notperceived at all by the human ear).

Each and every one of the rotor stages of the turbine that drives thefan via the gearbox may comprise in the range of from 60 to 140 rotorblades, for example in a range having a lower bound of any one of 70,75, 80, 85 or 90, and an upper bound of any one of 140, 130, 120 or 110,for example in the range of from 80 to 140 rotor blades.

The average number of rotor blades in a rotor stage of the turbine thatdrives the fan via the gearbox may be in the range of from 65 to 120rotor blades, for example in a range having a lower bound of any one of65, 70, 75, 80, 85 or 90, and an upper bound of any one of 120, 115, 110or 105, for example in the range of from 85 to 120 rotor blades.

The number of rotor blades in the most axially rearward turbine rotorstage of the turbine that drives the fan via the gearbox may be in therange of from 60 to 120 rotor blades, for example in a range having alower bound of any one of 60, 65, 70, 75, 80, 85 or 90, and an upperbound of any one of 120, 115, 110 or 105, for example 80 to 120 rotorblades.

In some arrangements, the contribution of the turbine to the EPNL at thetake-off lateral reference point may be in the range of from 15 EPNdBand 40 EPNdB lower than the contribution of the fan noise emanating fromthe rear of the engine to the EPNL at the take-off lateral referencepoint. In some arrangements, the contribution of the turbine to the EPNLat the take-off lateral reference point may be in the range of from 20EPNdB and 40 EPNdB, 25 EPNdB and 40 EPNdB, for example 25 EPNdB and 35EPNdB, for example 27 EPNdB and 33 EPNdB, lower than the contribution ofthe fan noise emanating from the rear of the engine to the EPNL at thetake-off lateral reference point.

In some arrangements, the ratio of the fan diameter to the diameter atthe leading edge of the tips of the most axially rearward turbine rotorstage of the turbine that drives the fan via the gearbox is in the rangeof from 2 to 3, optionally 2.3 to 2.9, optionally 2.4 to 2.8.

Arrangements of the present disclosure may be particularly beneficialfor fans that are driven via a gearbox. The input to the gearbox may bedirectly from the core shaft, or indirectly from the core shaft, forexample via a spur shaft and/or gear. The core shaft may rigidly connectthe turbine and the compressor, such that the turbine and compressorrotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only by the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox is a reduction gearbox (in that the output to the fan is alower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any reduction ratio (defined as the rotational speed ofthe input shaft divided by the rotational speed of the output shaft).For example, the gear ratio may be greater than 2.5 and/or less than 5.By way of more specific example, the gear ratio may be in the range offrom 3.2 to 5, or 3.2 or 3.3 to 4.2, or 3.3 or 3.4 to 3.7. By way offurther example, the gear ratio may be on the order of or at least 3,3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2, or betweenany two of the values in this paragraph. In some arrangements, the gearratio may be outside these ranges.

As noted elsewhere herein, the turbine that drives the fan via thegearbox may comprise at least two axially separated rotor stages. Theturbine that drives the fan via the gearbox may have a rotational speedat the take-off lateral reference point of Wlrp rpm. The minimum numberof rotor blades in any single rotor stage of the turbine that drives thefan via the gearbox may be given by NTURBmin. The diameter of the fanmay be given by ϕfan. A Low Speed System parameter (LSS) may be definedas:LSS=Wlrp×NTURBmin×ϕfan

The value of the Low Speed System parameter (LSS) may be given by:1.3×10⁶ m·rpm≤LSS≤2.9×10⁶ m·rpm

The value of the Low Speed System parameter (LSS) may be in a rangehaving a lower bound of any one of 1.3×10⁶ m·rpm, 1.4×10⁶ m·rpm, 1.5×10⁶m·rpm, 1.6×10⁶ m·rpm, 1.7×10⁶ m·rpm, 1.8×10⁶ m·rpm, or 1.9×10⁶ m·rpmand/or an upper bound of any one of 2.9×10⁶ m·rpm, 2.8×10⁶ m·rpm,2.7×10⁶ m·rpm, 2.6×10⁶ m·rpm, 2.5×10⁶ m·rpm, 2.4×10⁶ m·rpm, 2.3×10⁶m·rpm, or 2.2×10⁶ m·rpm.

Providing a gas turbine engine with a Low Speed System parameter (LSS)in the ranges described and/or claimed here has been found to result ina gas turbine engine that has high efficiency (for example due inparticular to high propulsive efficiency) and/or high thrust (forexample in the range of from 180 kN to 450 kN), but with acceptably low(and/or lower than conventional) noise levels (for example due toparticularly low turbine noise propagating from the rear of the engine).The rotational speed of the turbine that drives the fan via the gearboxat the take-off lateral reference point may be the same as, or similarto (for example within 5% of), the maximum rotational speed of thatturbine during take-off.

Purely by way of specific example, some gas turbine engines according tothe present disclosure may have a turbine rotational speed at thetake-off lateral reference point (Wlrp) in the range of from 5300 rpm to7000 rpm (for example 5700 rpm to 6500 rpm) and/or a fan diameter in therange of from 320 cm and 400 cm (for example 330 cm and 370 cm) and/or aminimum number of rotor blades in any single rotor stage of the turbinethat drives the fan via the gearbox (NTURBmin) in the range of from 70to 120 (for example 80 to 100).

Purely by way of further specific example, some gas turbine enginesaccording to the present disclosure may have a turbine rotational speedat the take-off lateral reference point (Wlrp) in the range of from 8000rpm to 9500 rpm (for example 8200 rpm to 9200 rpm) and/or a fan diameterin the range of from 220 cm and 290 cm (for example 230 cm and 270 cm)and/or a minimum number of rotor blades in any single rotor stage of theturbine that drives the fan via the gearbox (NTURBmin) in the range offrom 60 to 115 (for example 65 to 115, or 70 to 105).

The total number of turbine blades in the turbine that drives the fanvia the gearbox may be in the range of from 320 and 540, for example inthe range of from 330 to 500, or 340 to 450.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided downstream of the fan and compressor(s), forexample axially downstream. For example, the combustor may be directlydownstream of (for example at the exit of) the second compressor, wherea second compressor is provided. By way of further example, the flow atthe exit to the combustor may be provided to the inlet of the secondturbine, where a second turbine is provided. The combustor may beprovided upstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds).

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminum based material(such as an aluminum-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminum based body(such as an aluminum lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a schematic showing the measurement of Effective PerceivedNoise Level (EPNL) during take-off;

FIG. 5 is a graph showing an example of how the EPNL varies withdistance during take-off for an example of a gas turbine engine inaccordance with the present disclosure;

FIG. 6 is a graph showing the contribution of rear fan noise and forward(or front) fan noise to the EPNL at a take-off lateral reference pointfor an example of a gas turbine engine in accordance with the presentdisclosure;

FIG. 7 is a close-up schematic view of the turbine that drives the fanin an example of a gas turbine engine in accordance with the presentdisclosure; and

FIG. 8 is a diagram illustrating the calculation of fan tip relativeMach number.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Accordingly, the low pressure turbine 19 drivesthe fan 23 via the gearbox 30. Radially outwardly of the planet gears 32and intermeshing therewith is an annulus or ring gear 38 that iscoupled, via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

When in use to power an aircraft, the gas turbine engine 10 generatesnoise. As mentioned elsewhere herein, the gas turbine engine 10according to the present disclosure is arranged to reduce the noiseimpact whilst providing high efficiency.

Take-off is a particularly important flight condition from a noiseperspective, because the engine is typically being operated at a highpower condition, and because the engine is close to the ground, and thuspotentially close to communities. In order to quantify the impact of thegenerated noise as perceived by the human ear, an “Effective PerceivedNoise Level” (EPNL) is defined. The EPNL takes into account factors suchas frequency, absolute level, tonal components and duration of thenoise, and is calculated in the manner defined in Appendix 2 of theFifth Edition (July 2008) of Annex 16 (Environmental Protection) to theConvention on International Civil Aviation, Volume 1 (Aircraft Noise).

A take-off lateral reference point is used in order to quantify theimpact of the generated noise specifically during take-off of anaircraft powered by the gas turbine engine 10, as defined in Section3.3.1, a), 1) of the Fifth Edition (July 2008) of Annex 16(Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise).

In particular, the take-off lateral reference point is defined as thepoint on a line parallel to and 450 m from the runway centre line wherethe EPNL is a maximum during take-off. This is illustrated in FIG. 4. Inparticular, FIG. 4 shows a series of noise-measurement devices 150, suchas microphones, positioned along a line A on the ground that is 450 mfrom the take-off path (which may be referred to as the runwaycentreline) of an aircraft 100 powered by one or more (for example 2 or4) gas turbine engines 10. Each microphone 150 measures the noise at itsrespective location during take-off, and the measurements are used tocalculate the EPNL at that location. In this way, it is possible todetermine the EPNL along the line A (450 m from the runway centreline,extended forwards along the ground after lift-off).

FIG. 5 shows an example of a graph showing EPNL in dB (EPNdB) along theline A against the distance from the start of take-off (which may bereferred to as distance from release of brake, indicating that it is thedistance from the point at which the aircraft starts its main take-offacceleration at the start of the runway). As illustrated, the EPNL ofthe engine initially increases, and this increase may continue evenafter lift-off (i.e. after the point at which the aircraft loses contactwith the ground), which is labelled as the point “LO” in FIG. 5, purelyby way of example.

At a certain position on the flight path, the EPNL (i.e. the EPNL asmeasured on the ground, along line A in FIG. 4) reaches a maximum, andthen starts to fall. The distance along line A (i.e. the distance alongthe runway centreline) at which this occurs is the take-off lateralreference point (labelled RP in FIG. 5). The EPNL at the take-offlateral reference point RP (labelled RP EPNL in FIG. 5) is the maximumEPNL during take-off.

The take-off period may be considered to last at least as long asnecessary to determine the maximum point (at distance RP) of the EPNLbetween release of brake and top of climb of the aircraft. In practice,this is likely to be within a horizontal distance of 10 km or less ofthe release of brake.

A number of different noise sources contribute to the EPNL, and thus tothe RP EPNL. In a conventional engine, noise from the fan that isemitted from the front of the engine and the rear of the engine producesa significant contribution to the RP EPNL.

The present inventors have found that the contribution to the RP EPNL ofthe noise emitted from the front of the engine in particular can besignificantly reduced through the appropriate use of a gearbox 30through which the fan 23 is driven by the turbine 19.

FIG. 6 shows the total engine EPNL at the take-off lateral referencepoint (RP EPNL), along with the contributions to the total engine EPNLat that point of both the rear fan noise and the front fan noise for anexample of a gas turbine engine 10 in accordance with the presentdisclosure. Referring back to FIG. 1, the front fan noise is illustratedby arrow V as the noise from the fan emanating from the front of theengine, and the rear fan noise is illustrated by arrow U as the noisefrom the fan emanating from the rear of the engine. The contribution ofthe front fan noise is X EPNdB lower than the rear fan noisecontribution, where X is in the range of from 0 EPNdB to 12 EPNdB.Purely by way of non-limitative example, the value of X for the gasturbine engine 10 having the noise characteristic illustrated in FIG. 6may be on the order of 8 EPNdB.

Accordingly, the gas turbine engine 10 according to the presentdisclosure can be particularly efficient—for example having highpropulsive efficiency through having the fan 23 driven via a gearbox30—whilst having reduced noise signature due to the relative reductionin noise (as measured by EPNL) of the fan emanating from the front ofthe engine. Of course, the total engine noise comprises other noisesources in addition to the fan noise, such as (by way of non-limitativeexample) turbine noise and jet noise.

It will be appreciated that the individual contributions of thecomponents (such as the noise from the fan 23 that emanates from therear of the engine, the noise from the fan 23 that emanates from thefront of the engine, and the noise from the jet and the noise from theturbine 19) can be identified through conventional analysis of the noisemeasured by the microphones 150. For example, each component has afrequency signature that can be predicted, meaning that noise that isgenerated in accordance with the predicted frequency signature can beattributed to that component. In practice, the noise that is generatedby the fan and emanates from the rear of the engine may be distinguishedfrom the noise that is generated by the fan and emanates from the frontof the engine using a source location technique, such as measuring thephase difference of the noise. In this regard, the noise that isgenerated by the fan and emanates from the rear of the engine isphase-shifted relative to the noise that is generated by the fan andemanates from the front of the engine due to the physical separation ofthe front and rear of the engine.

Returning to FIG. 1, the flow exiting from the bypass duct 22 isillustrated by arrow P. This flow may be a significant contributor (forexample the major contributor) to the jet noise, for example due to theflow P generating a shear layer S with the surrounding air as it exitsthe engine. The average velocity of the flow P at the exit to the bypassduct may be within the ranges described and/or claimed herein (forexample in the range of from 200 m/s to 275 m/s) at the take-off lateralreference condition, thereby reducing noise at that condition.Additionally or alternatively, the average velocity of the flow P at theexit to the bypass duct at the take-off lateral reference point may bein the range of from 50 m/s to 90 m/s lower, optionally 55 m/s to 85 m/slower, optionally 60 m/s to 85 m/s lower than the average velocity ofthe flow at the exit to the bypass duct at cruise conditions, therebyreducing noise at the take-off lateral reference point.

The average velocity of the flow P at the exit to the bypass duct 22 maybe the mass-averaged flow velocity at the exit plane Z that isperpendicular to the engine axis 9 and passes through the trailing edge210 of the nacelle 21.

FIG. 1 also shows noise attenuation material 310 provided to an intakeof the engine, noise attenuation material 320 provided to the outer flowboundary of the bypass duct, and noise attenuation material 330 providedto the inner flow boundary of the bypass duct.

A bypass noise attenuation proportion L may be defined as:

$L = \frac{J + H}{2G}$

where:

G is the axial length between the tip of the trailing edges of the fanblades and the trailing edge of the nacelle;

H is the total axial length of acoustic attenuation material provided tothe outer flow boundary of the bypass duct over the axial extent betweenthe tip of the trailing edges of the fan blades and the trailing edge ofthe nacelle; and

J is the total axial length of acoustic attenuation material provided tothe inner flow boundary of the bypass duct over the axial extent betweenthe tip of the trailing edges of the fan blades and the trailing edge ofthe nacelle.

An intake noise attenuation proportion k may be defined as:

$K = \frac{E}{F}$

where:

E is the total axial length of acoustic attenuation material provided tothe intake; and

F is the axial length of the intake.

Examples of the lengths E, F, G, H and J are shown in FIG. 1. It will beappreciated that the total axial length of acoustic attenuation material(i.e. the values F, H, and J) may be provided as a single segment (as inthe case for F and J in the illustrated example), or in multiplesegments (as in the case for H in the illustrated example, whichcomprises a first segment H1 and a second segment H2, with the value ofH being given by H=H1+H2).

A forward to rearward noise attenuation proportion M may be in theranges described and/or claimed here, for example from 0.8 to 2.5, 1.1to 2.3, 1.2 to 2.1, 1.3 to 2, 1.4 to 1.8, or on the order of 1.6, where:

$M = \frac{K}{L}$

The bypass noise attenuation proportion L may be in the ranges describedand/or claimed herein, for example from 0.4 to 0.7, optionally, 0.45 to0.65, optionally 0.5 to 0.6. The values of H/G and J/G may be withinthese ranges individually.

The intake noise attenuation proportion K may be in the ranges describedand/or claimed herein, for example from 0.55 to 0.95, optionally, 0.6 to0.9, optionally 0.7 to 0.8.

The acoustic attenuation material may take any suitable form, forexample as described elsewhere herein.

A further example of a feature that may be better optimized for gasturbine engines 10 according to the present disclosure compared withconventional gas turbine engines is the intake region, for example theratio between the intake length L and the fan diameter D. Referring toFIG. 1, the intake length L is defined as the axial distance between theleading edge of the intake and the leading edge of the root of the fanblade, and the diameter D of the fan 23 is defined at the leading edgeof the fan 23. Gas turbine engines 10 according to the presentdisclosure, such as that shown by way of example in FIG. 1, may havevalues of the ratio L/D as defined herein, for example less than orequal to 0.5, for example in the range of from 0.25 to 0.5, optionally0.3 to 0.5, optionally 0.32 to 0.49, optionally 0.33 to 0.48. This maylead to further advantages, such as installation and/or aerodynamicbenefits, whilst maintaining forward fan noise at an acceptable level.

FIG. 8 illustrates a view onto the radially outermost tip of one of thefan blades of the fan 23. In use, the fan blade rotates, such that thetip has a rotational velocity given by the rotational speed of the fanmultiplied by the radius of the tip. The rotational velocity at theleading edge of the tip (i.e. using the radius of the leading edge ofthe tip) can be used to calculate the rotational Mach number at the tip,illustrated by Mn_(rot) in FIG. 8.

The axial Mach number at the leading edge of the tip of the fan blade isillustrated as Mn_(axial) in FIG. 8. In practice (and as used tocalculate the fan tip relative Mach number Mn_(rel) as used herein),this may be approximated by multiplying the average axial Mach numberover the plane that is perpendicular to the axial direction at theleading edge of the tip of the fan blade by 0.9.

The fan tip relative Mach number (Mn_(rel)) is calculated as the vectorsum of the axial Mach number Mn_(axial) and the rotational Mach numberat the tip Mn_(rot), i.e. having a magnitude Mnrel=√{square root over(Mnaxial²+Mnrot²)}.

In order to calculate the Mach numbers (Mn_(axial) and Mn_(rot)) fromthe velocities, the average static temperature over the plane that isperpendicular to the axial direction at the leading edge of the tip ofthe fan blade is used to calculate the speed of sound.

The fan tip relative Mach number (Mn_(rel)) may be in the rangesdescribed and/or claimed herein, for example no greater than 1.09 and/orin the range of from 0.8 M to 1.09 M, optionally 0.9 M to 1.08 M,optionally 1.0 M to 1.07 M at the take-off lateral reference point.

Accordingly, the fan noise, including at least the noise propagatingfrom the front of the engine at the take-off lateral reference point,may be reduced compared with engines of comparable size and/or power.Additionally or alternatively, the reduced fan tip relative Mach numbermay at least in part facilitate a lower jet velocity, which may in turnlead to lower jet noise.

As mentioned elsewhere herein, a number of different noise sourcescontribute to the EPNL, and thus to the RP EPNL. In a conventionalengine, the turbine that drives the fan also provides a significantcontribution to the RP EPNL.

However, the contribution to the RP EPNL of the turbine 19 that drivesthe fan 23 via the gearbox 30 may also be reduced by increasing thefrequencies of the fundamental tones generated by the turbine tofrequencies that are less well perceived by the human ear and/or haveincreased atmospheric attenuation, thereby reducing the perceived noisefrequency rating. As such, these tones are given a lower weighting inthe EPNL calculation (even a zero weighting if the frequency is highenough), thereby reducing the contribution of the turbine noise to theRP EPNL.

Reducing the turbine noise may be particularly beneficial for engineshaving reduced jet noise and/or reduced forward fan noise, because thecontribution of the turbine noise to the overall noise of the enginewould otherwise be more significant due to the reduced fan/jet exitnoise.

FIG. 7 shows the turbine 19 that drives the fan 23 via the gearbox 30 inmore detail for the gas turbine engine 10 according to an example of thepresent disclosure, which may be referred to as the low pressure turbine19. The low pressure turbine 19 comprises four rotor stages 210, 220,230, 240. The low pressure turbine 19 is therefore a four stage turbine19. However, it will be appreciated that the low pressure turbine 19 mayconsist of other numbers of turbine stages, for example three or five.

Each rotor stage 210, 220, 230, 240 comprises rotor blades that extendbetween an inner flow boundary 250 and an outer flow boundary 260. Eachof the rotor stages 210, 220, 230, 240 is connected to the same coreshaft 26 that provides input to the gearbox 30. Accordingly, all of therotor stages 210, 220, 230, 240 rotate at the same rotational speed Wlaround the axis 9 in use. In the FIG. 7 example the rotor stages 210,220, 230, 240 each comprise a respective disc 212, 222, 232, 242supporting the rotor blades. However, it will be appreciated that insome arrangements the disc may not be present, such that the blades aresupported on a circumferentially extending disc.

Each rotor stage 210, 220, 230, 240 has an associated stator vane stage214, 224, 234, 244. In use, the stator vane stages do not rotate aroundthe axis 9. Together, a rotor stage 210, 220, 230, 240 and itsassociated stator vane stage 214, 224, 234, 244 may be said to form aturbine stage.

The lowest pressure rotor stage 210 is the most downstream rotor stage.The rotor blades of the lowest pressure rotor stage 210 are longer (i.e.have a greater span) than the rotor blades of the other stages 220, 230,240. Indeed, each rotor stage has blades having a span that is greaterthan the blades of the upstream rotor stages.

The number of rotor blades may have an impact on the frequency of thesound generated by the turbine 19. The rotational speed Wl of the lowpressure turbine 19 may also have an effect on the frequency of thesound generated by the turbine 19, and this, in turn, is linked to therotational speed of the fan 23 by the gear ratio of the gearbox 30.

Each rotor stage 210, 220, 230, 240 consists of any desired number ofrotor blades. For example, each and every one of the rotor stages 210,220, 230, 240 of the turbine 19 that drives the fan 23 via the gearbox30 may comprise in the range of from 80 to 140 rotor blades. By way offurther example, the average number of rotor blades in a rotor stage210, 220, 230, 240 of the turbine 19 that drives the fan 23 via thegearbox 30 may be in the range of from 85 to 120 rotor blades. By way offurther example, the number of rotor blades in the most axially rearwardturbine rotor stage 210 of the turbine 19 that drives the fan 23 via thegearbox 30 may be in the range of from 80 to 120 rotor blades.

In one particular, non-limitative example, the first (most upstream)rotor stage 240 and the second rotor stage 230 may each comprise around100 rotor blades, and the third rotor stage 220 and fourth (mostdownstream) rotor stage 210 may each comprise around 90 rotor blades.However, it will be appreciated that this is purely by way of example,and the gas turbine engine 10 in accordance with the present disclosuremay comprise other numbers of turbine blades, for example in the rangesdefined elsewhere herein.

At the take-off lateral reference point, the low pressure turbine 19 hasa rotational speed of Wlrp rpm. In one example, the low pressure turbine19 of gas turbine engine 10 has a rotational speed at the take-offlateral reference point in the range of from 5300 rpm to 7000 rpm. Inthis example, the diameter of the fan 23 (as defined elsewhere herein)may be in the range of from 320 cm to 400 cm. In one specific,non-limitative example, the low pressure turbine 19 of the gas turbineengine 10 has a rotational speed at the take-off lateral reference pointof around 5900 rpm, and a fan diameter of around 340 cm.

In one example, the low pressure turbine 19 of gas turbine engine 10 hasa rotational speed at the take-off lateral reference point in the rangeof from 8000 rpm to 9500 rpm. In this example, the diameter of the fan23 (as defined elsewhere herein) may be in the range of from 220 cm to290 cm. In one specific, non-limitative example, the low pressureturbine 19 of the gas turbine engine 10 has a rotational speed at thetake-off lateral reference point of around 8700 rpm, and a fan diameterof around 240 cm.

A Low Speed System parameter (LSS) may be defined for the gas turbineengine 10 as:LSS=Wlrp×NTURBmin×ϕfan

where:

Wlrp is the rotational speed at the take-off lateral reference point ofthe turbine 19 that drives the fan 23 via the gearbox 30 (rpm);

NTURBmin is the minimum number of rotor blades in any single rotor stage210, 220, 230, 240 of the turbine 19 that drives the fan 23 via thegearbox 30; and

ϕfan is the diameter of the fan (m).

In some arrangements, the Low Speed System parameter (LSS) for the gasturbine engine 10 is in the range:1.3×10⁶ m·rpm≤LSS≤2.9×10⁶ m·rpm

Purely by way of non-limitative example, the gas turbine engine 10 mayhave a fan diameter of 3.4 m, a minimum number of rotor blades in anysingle rotor stage 210, 220, 230, 240 of 100, and a rotational speed atthe take-off lateral reference point of the low pressure turbine 19 of5900 rpm, giving a Low Speed System parameter (LSS) of around 2.0×10⁶.

Purely by way of further non-limitative example, the gas turbine engine10 may have a fan diameter of 2.4 m, a minimum number of rotor blades inany single rotor stage 210, 220, 230, 240 of 95, and a rotational speedat the take-off lateral reference point of the low pressure turbine 19of 8700 rpm, giving a Low Speed System parameter (LSS) of around2.0×10⁶.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

The invention claimed is:
 1. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein: a contribution of fan noise emanating from a front of the gas turbine engine to an Effective Perceived Noise Level (“EPNL”) at a take-off lateral reference point, defined as a point on a line parallel to and 450 m from a runway centre line where the EPNL is a maximum during take-off, is in a range of from 0 EPNdB to 12 EPNdB lower than a contribution of fan noise emanating from a rear of the gas turbine engine to the EPNL at the take-off lateral reference point: a bypass duct is defined between an inner flow boundary formed by the engine core and an outer flow boundary formed by a nacelle; an intake is defined as a radially outer flow boundary of a flow into the gas turbine engine upstream of tips of the leading edge of the plurality of fan blades; a bypass noise attenuation proportion L is defined as: $L = {\frac{J + H}{2G}.}$ where: G is an axial length between tips of the trailing edges of the plurality of fan blades and a trailing edge of the nacelle; H is a total axial length of acoustic attenuation material provided to the outer flow boundary of the bypass duet over an axial extent between the tips of the trailing edges of the plurality of fan blades and the trailing edge of the nacelle; and J is a total axial length of acoustic attenuation material provided to the inner flow boundary of the bypass duct over the axial extent between the tips of the trailing edges of the plurality of fan blades and the trailing edge of the nacelle; an intake noise attenuation proportion k is defined as: $K = {\frac{E}{F}.}$ where: E is the total axial length of acoustic attenuation material provided to the intake; and F is the axial length of the intake and a forward to rearward noise attenuation proportion M is in a range of from 0.8 to 2.5, where: $M = {\frac{K}{L}.}$
 2. The gas turbine engine according to claim 1, wherein the contribution of the fan noise emanating from the front of the engine to the EPNL at the take-off lateral reference point is in range of from 0 EPNdB to 1.0 EPNdB lower than the contribution of the fan noise emanating from the rear of the gas turbine engine to the EPNL at the take-off lateral reference point.
 3. The gas turbine engine according to claim 1, wherein a relative Mach number at a tip of each fan blade of the plurality of fan blades does not exceed 1.09 M at the take-off lateral reference point.
 4. The gas turbine engine according to claim 1, wherein the intake noise attenuation proportion K divided by the bypass noise attenuation proportion L is in a range of from 1.1 to 2.3.
 5. The gas turbine engine according to claim 1, wherein the bypass noise attenuation proportion L is in a range of from 0.4 to 0.7.
 6. The gas turbine engine according to claim 1, wherein the intake noise attenuation proportion K is in a range of from 0.55 to 0.95.
 7. The gas turbine engine according to claim 1, wherein a gear ratio of the gearbox is in a range of from 3 to
 5. 8. The gas turbine engine according to claim 1, wherein: a fan diameter is in a range of from 320 cm to 100 cm and, when the EPNL at the take-off lateral reference point is a maximum during take-off, the rotational speed of the fan is in a range of from 1300 rpm to 1800 rpm; or the fan diameter is in a range of from 220 cm to 290 cm, and when the EPNL at the take-off lateral reference point is a maximum during take-off, the rotational speed of the fan at the take-off lateral reference point is in a range of from 2000 rpm to 2800 rpm.
 9. The gas turbine engine according to claim 1, wherein, when the EPNL at the take-off lateral reference point is a maximum during take-off, a fan tip pressure ratio, defined as a ratio of a mean total pressure of a low at a fan exit that subsequently flows through the bypass duct to a mean total pressure of a flow at a fan inlet, is in a range of from 1.25 to 1.5.
 10. The gas turbine engine according to claim 1, wherein a contribution of the turbine to the EMIL at the take-off lateral reference point is in a range of from 15 EPNdB to 40 EPNdB, lower than the contribution of the fan noise emanating from the rear of the engine to the EPNL at the take-off lateral reference point.
 11. The gas turbine engine according to claim 1, wherein: the turbine comprises at least two axially separated rotor stages; and each and every one of the rotor stages of the turbine comprises in a range of from 60 to 140 rotor blades.
 12. The gas turbine engine according to claim 1, wherein: the turbine comprises at least two axially separated rotor stages; and an average number of rotor blades in a rotor stage of at least two axially separated rotor stages is in a range of from 65 to 120 rotor blades.
 13. The gas turbine engine according to claim 1, wherein: the turbine comprises at least two axially separated rotor stages; and a number of rotor blades in a most axially rearward turbine rotor stage of the at least two axially separated rotor stages is in a range of from 60 to 120 rotor blades.
 14. The gas turbine engine according to claim 1, wherein: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.
 15. The gas turbine engine according to claim 1, further comprising an intake that extends upstream of the plurality of fan blades, wherein: a radially inner intake length L is defined as an axial distance between a leading edge of the intake and the leading edges of the plurality of fan blades at a hub; a fan diameter D is a diameter of the fan at a leading edges of tips of the plurality of fan blades; and a ratio LID is in a range of from 0.2 to 0.5.
 16. An aircraft comprising the gas turbine engine according to claim
 1. 17. A method of operating an aircraft comprising a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, and the method comprising taking off from a runway, during which: a contribution of the fan noise emanating from a front of the gas turbine engine to an Effective Perceived Noise Level (“EPNL”) at a take-off lateral reference point, defined as a point on a line parallel to and 450 m from a runway centre line where the EPNL is a maximum during take-off, is in a range of from 0 EPNdB to 12 EPNdB, lower than the contribution of fan noise emanating from a rear of the gas turbine engine to the EPNL, at the take-off lateral reference point, wherein: a bypass duct is defined between an inner flow boundary formed by the engine core and an outer flow boundary formed by a nacelle; an intake is defined as a radially outer flow boundary of a flow into the gas turbine engine upstream of tips of the leading edge of the plurality of fan blades; a bypass noise attenuation proportion L is defined as: $L = {\frac{J + H}{2G}.}$ where: G is an axial length between tips of trailing edges of the plurality of fan blades and a trailing edge of the nacelle; H is a total axial length of acoustic attenuation material provided to the outer flow boundary of the bypass duct over an axial extent between the tips of the trailing edges of the plurality of fan blades and the trailing edge of the nacelle; and J is a total axial length of acoustic attenuation material provided to the inner flow boundary of the bypass duet over the axial extent between the tips of the trailing edges of the plurality of fan blades and the trailing edge of the nacelle; an intake noise attenuation proportion L is defined as: $K = {\frac{E}{F}.}$ where: E is the total axial length of acoustic attenuation material provided to the intake; and F is the axial length of the intake; and a forward to rearward noise attenuation proportion NI is in a range of from 0.8 to 2.5, where: $M = {\frac{K}{L}.}$
 18. The method according to claim 17, wherein: the bypass noise attenuation proportion L is in a range of from 0.4 to 0.7; and/or the intake noise attenuation proportion K is in a range of from 0.55 to 0.95. 